A satellite is placed in orbit through the combination of a launcher space vehicle and of its own propulsion systems. The launcher transports and releases the satellite on a first, so-called transfer, terrestrial orbit, whose perigee is generally low; once on this first orbit, a propulsion system of the satellite takes over so as to transport the satellite to its final orbit. Generally, this transfer is carried out by means of a principal satellite propulsion unit PSP consuming a chemical fuel of propellant type, delivering a high-power thrust making it possible to quickly reach the final orbit.
Once in post, several secondary propulsion units of lower powers maintain the satellite in position on the orbit. Accordingly, propellant-based chemical propulsion units or electric propulsion units can be used. In an electric propulsion unit, of the plasma propulsion unit or ion propulsion unit type, xenon atoms are ionized by collision with electrons, creating xenon ions. Thrust is generated when the charged xenon ions are accelerated out of the propulsion unit by an electromagnetic field. Although expensive and of significant initial mass, the effectiveness of the propulsion unit, or its capacity to generate force by ejection of mass, also called specific impulse, is substantially more significant than that of chemical propulsion units.
In the known systems, chemical propulsion units and electric propulsion units are positioned at several locations on the structure of the satellite to meet all the requirements of the mission, from transport from the transfer orbit to maintaining in orbit throughout the life of the satellite. The drawback of the propulsion systems thus implemented is a high cost and high mass, of the various propulsion units and of the fuel. These drawbacks limit the satellite's payload carrying capacity.
According to the known state of the art, an orbit control system seeks to pilot the position of the satellite through six orbital parameters. FIG. 1 represents a geostationary satellite 10 in orbit 11 around the earth 12. The orbit 11 is inclined by an angle θ with respect to the equatorial plane 13 which contains the ideal geostationary orbit 14. The orbit 11 of the satellite cuts the equatorial plane 13 at two points 15 and 16, commonly called orbital nodes. The six orbital parameters used to convey the position of a satellite are also known: the semi-major axis, the eccentricity, the inclination, the argument of the ascending node, the argument of the perigee, and the true anomaly. Orbit control consists in quantifying these orbital parameters and in carrying out the operations required by means of the on-board propulsion systems, to maintain the satellite in a predefined zone around an ideal position. By way of example, for a geostationary satellite, a drift window of plus or minus 0.1°, representing a width of almost 150 km, is allocated around a target position.
A contemporary architecture, such as represented in FIG. 2, of a satellite 10 comprises a parallelepipedal structure 20 on which are fixed various devices useful for the piloting of the satellite 10 and for its mission. Telecommunications instruments 21 are installed on a face 22 whose orientation is maintained towards the earth, commonly called the earth face. On an opposite face 23, commonly called the anti-earth face, is positioned the principal satellite propulsion unit PSP which ensures notably the thrust necessary for transfer from the low orbit to the final orbit. On two opposite lateral faces 24 and 25, commonly called the north face and the south face, because of their orientation with respect to the equatorial plane, are positioned two suites of solar panels 26 and 27 allowing electrical energy supply to the onboard systems. Various devices can be on board the lateral faces 28 and 29, commonly called the east and west face for their orientation with respect to a terrestrial longitude. The maintaining of a constant orientation of the satellite with respect to the earth is necessary for the proper conduct of the mission of the satellite, for example to orient the solar panels 26 and 27 or to point the telecommunications systems 21 towards the earth. This is carried out by means of an attitude control system. Several attitude control systems able to detect and correct orientation errors are known. Thus, the measurement of the orientation of the satellite can be carried out by means of a sensor set, comprising for example a sensor directed towards the earth, positioned on the earth face for a measurement on two axes, pitch and roll, with respect to the earth and a set 30 of gyroscopes for detecting the rotation rates in relation to three axes. On the basis of these measurements, corrections of orientation of the satellite about its centre of gravity can be carried out, for example by means of a set of inertia wheels 31 or of gyroscopic actuators.
A satellite equipped with such a system allowing attitude control is said to be stabilized in relation to three axes. Typically, by controlling the rotation rate and the orientation of the inertia wheels, it is possible to correct an orientation error in a reference trihedron tied to the satellite. Hereinafter, an axis directed towards the earth is called Z, also called the yaw axis, an axis perpendicular to the orbit and oriented in the opposite sense to the angular momentum of the orbit (towards the south for a geostationary) is called Y, also called the pitch axis, and an axis forming with Y and Z a right-handed orthogonal reference frame is called X, also called the roll axis which is oriented along the velocity in the case of circular orbits.
For orbit control, several propulsion units are disposed on the structure 20 of the satellite 10. A first propulsion unit of high power PSP, making it possible to ensure transfer between the initial terrestrial orbit (after launcher release) and the final orbit, is positioned on the anti-earth face 23. According to a known state of the art, a first set of propulsion units, comprising for example two propulsion units 32 and 33 positioned at the north face and at the south face in proximity to the anti-earth face, is used to control the inclination. A second set of propulsion units, such as for example the propulsion units 34 and 35, positioned at the east and west faces, is used for the control of the eccentricity and the drift. It is also known that control of the inclination requires of the order of five to ten times more fuel than control of the eccentricity and of the drift. For this reason, inclination control is in general carried out by means of plasma propulsion unit, consuming less fuel, while the propulsion units dedicated to the control of the eccentricity and of the drift are usually based on chemical propellant.
By way of example, a contemporary satellite of dry mass 2500 kg and making it possible to carry a payload of 900 kg, comprises a principal propulsion unit, two plasma propulsion units for the inclination and the eccentricity, and four propellant-based propulsion units for the eccentricity and the drift. Typically, 1700 kg of propellant are necessary for the initial orbit transfer, and 220 kg of Xenon are necessary to ensure orbit control of the satellite for a mission duration of about 15 years. Thus, the cost and mass of current propulsion systems limit the capacity to carry a high payload. Note also that in most known propulsion systems for orbit control, the various onboard propulsion units comprise in reality two propulsive motors positioned side by side, for mission safety and reliability reasons. This redundancy, well known to the person skilled in the art, is not represented in the figures but it is considered hereinafter that a propulsion unit can consist of one or more propulsive motors forming a propulsive assembly, and whose deliverable thrust is identical, in orientation or in intensity.
FIGS. 3a, 3b and 3c illustrate the principle of orbit control for a satellite according to the known state of the art. The structure 20 of the satellite 10 is represented in a side view, the east face being visible. The propulsion unit 32 is connected to the north face of the structure 20 by means of a two-axis mechanism 40. The two-axis mechanism 40 allows the rotation of the propulsion unit 32 with respect to the structure 20 in relation to a first axis parallel to the Y axis and a second axis parallel to the X axis. In FIGS. 3a to 3c, the two-axis mechanism 40 is a gimbal link produced by means of a first pivot link 41 with axis parallel to the Y axis and a second pivot link 42 with axis parallel to the X axis. The centre of mass of the satellite, situated inside the parallelepipedal structure 20, is referenced CM.
In FIG. 3a, the orientation of the propulsion unit 32 makes it possible to direct the thrust of the propulsion unit towards the centre of mass CM of the satellite. To perform an inclination correction manoeuvre, a technique known to the person skilled in the art consists in firing the propulsion unit 32 a first time in proximity to an orbital node, for example 15, and then the propulsion unit on the opposite side a second time in proximity to the opposite orbital node, 16 in the example. Thus, the thrust, oriented towards the centre of mass CM, of the first firing of the propulsion unit 32 displaces the satellite in a direction having a Z component and a Y component. Twelve hours later, the thrust of the second firing at the opposite orbital node, displaces the satellite in a direction having an opposite Z component to the first firing, and which compensates for the undesired effect thereof on the eccentricity and a likewise opposite Y component but whose desired inclination effects are aggregated. Thus, two firings of equal intensities carried out at twelve hourly intervals in proximity to the orbital nodes 15 and 16 make it possible to cancel the effect of the radial component and thus to preserve only a north-south correction. This known procedure allows daily correction of the inclination.
Through this same technique it is also possible, by applying a second thrust of different intensity from the first, to apply eccentricity corrections in relation to an axis perpendicular to the line joining the two orbital nodes 15 and 16. Techniques have also developed for allowing eccentricity corrections in relation to a second axis, by offsetting the firing of the propulsion unit with respect to the orbital node, but at the price of less effective control of the inclination. To summarize, the known systems make it possible by means of two propulsion unit systems 32 and 33 to ensure control of the inclination and control of the eccentricity in relation to an axis without deoptimization of the inclination control, or to ensure control of the inclination and control of the eccentricity in relation to two axes with deoptimization of the inclination control. Control of the drift may not be carried out by these two propulsion units. A contemporary satellite accordingly comprises four chemical-propellant nozzles positioned on the east and west faces of the satellite.
The propulsion unit systems 32 and 33 are also useful for managing the momentum of the attitude control systems, as illustrated in FIGS. 3b and 3c. By applying a thrust outside of the centre of mass CM—in a plane Y-Z in FIG. 3b and outside of the plane Y-Z in FIG. 3c, a rotation couple is generated on the satellite—a roll couple in FIG. 3b and a pitch and yaw couple in FIG. 3c. These two couples can be used to load or unload the inertia wheels in relation to two axes. For example, when the rotation rate of an inertia wheel reaches its limit velocity, it will be sought to orient the thrust intentionally outside of the centre of mass CM so as to generate, in addition to the desired displacement of the satellite, a couple making it possible to desaturate the inertia wheel, or more generally, the problem will be anticipated by reducing the angular momentum to desired values upon each manoeuvre. These desired values being able of course to be zero, but also a value judiciously defined so as to anticipate the evolution of the angular momentum between two manoeuvres under the effect of the solar radiation pressure notably.
Note also that the centre of mass of the satellite varies in the course of the life of the satellite, notably because of the progressive consumption of the onboard fuel. In the known systems, algorithms are implemented for the combined management of the attitude control and of the orbit control, and to make it possible to take into account the position of the centre of mass CM throughout the life of the satellite.
The issue of being able to have effective propulsion systems is therefore understood. The current solutions, which implement different nature propulsion units at various locations of the satellite are relatively complex and expensive and exhibit a high mass which limits the satellite's carrying capacity.